This invention relates generally to the field of combustion turbine engines and more particularly to the use of ceramic matrix composite materials in a combustion turbine engine.
U.S. Pat. No. 6,197,424 describes a ceramic insulating material that may be applied to a ceramic matrix composite (CMC) material for use in high temperature applications such as a gas turbine engine. That patent illustrates several components of a gas turbine engine utilizing the insulated CMC material, however, that patent does not describe how the insulated CMC material may be secured to the metal casing of the gas turbine engine.
U.S. Pat. No. 4,759,687 illustrates the use of a ceramic composition for a turbine ring application. The method of attachment described in this patent disadvantageously results in portions of the metal structure of the turbine ring remaining exposed to the hot combustion gasses.
Ceramic coatings are often applied directly to metal components to increase the high temperature performance characteristics of the components. The differential thermal expansion characteristics of metal and ceramic presents a design challenge for such coatings, as discussed in U.S. Pat. No. 5,080,557.
U.S. Pat. No. 4,679,981 describes an arrangement for clamping an abradable ceramic turbine blade ring so that there is always a compressive force on the ring. This arrangement relies on the differential cooling of the underlying metal carrier and it purposefully provides no cooling for the ceramic material. The safe operating temperature of the ceramic material would thus limit applications of this design.
U.S. Pat. No. 5,363,643 describes a ceramic combustor liner for a gas turbine engine. A plurality of individual ceramic liner segments is rigidly attached with a bolt and nut combination to an outer frame to form the cylindrical combustor shape. Each liner segment is carried by the outer frame and moves therewith as the frame expands and contracts, thereby mitigating the stresses experienced by the individual segments. This design necessitates the use of a large number of individual segments, which in turn results in a large number of joints where leakage of cooling air may occur. Such air leakage has a detrimental impact on engine efficiency and should be minimized. Furthermore, the use of small fasteners inside a gas turbine engine is generally undesirable.
U.S. Pat. No. 4,907,411 describes the use of sheet metal mounting members to support ceramic combustion chamber segments. The sheet metal members are used to space the ceramic segments relative to a housing, but they offer no structural support for the ceramic segments. As such, this attachment arrangement would be of limited value in applications where mechanical loads may be imposed upon the ceramic material, such as in a turbine shroud ring application where a ceramic ring segment may be exposed to impact with rotating turbine blades. Furthermore, this design requires the placement of a thermally insulating material between the sheet metal members and the ceramic combustion chamber segments. The ceramic material in this design is a non-oxide material such as silicon carbide or silicon nitride that is relatively very conductive to heat (10-20 watts/meter-xc2x0 K). This design allows the ceramic material to operate at a high temperature, and it provides protection to the metal members through the use of the insulating sealing strip between the metal and the ceramic, a layer of thermally reflective material on the side of the metal that faces the ceramic, and a small flow of cooling fluid between the metal and the ceramic surfaces.
Thus, improved manners of attaching a ceramic matrix composite material to a turbine casing are needed to provide thermal protection to metal parts, to eliminate the need for small fasteners and intervening insulating members, and to provide mechanical support for applications where mechanical loads are imposed onto the CMC material.
A component for use in a combustion turbine engine is described herein as including: a metal support member supported within a casing of a gas turbine engine and further comprising an extending portion; a ceramic matrix composite member shielding the metal support member from a combustion gas flowing within the combustion turbine engine during operation of the combustion turbine engine and comprising an arcuate portion extending around and in direct contact with the extending portion of the metal support member for supporting the ceramic matrix composite member from the metal support member; and the ceramic matrix composite member selected to have a thermal conductivity characteristic that is sufficiently low to maintain the support member below a predetermined temperature during operation of the combustion turbine engine. The ceramic matrix composite member may be separated from the metal support member by a gap having a predetermined maximum dimension at a location remote from the arcuate portion, the predetermined maximum dimension selected to control a level of stress developed in the shroud member when the ceramic matrix composite member is deflected to reduce the gap to zero.
A blade shroud assembly for a combustion turbine engine is described herein as including: a metal support member supported within a combustion turbine engine and comprising an upstream edge and an opposed downstream edge each extending along a circumferential length; a ceramic matrix composite shroud member comprising an upstream portion and an opposed downstream portion each extending along a circumferential length and each having an arcuate shape defining an upstream slot and a downstream slot receiving and in direct contact with respectively the upstream edge and the downstream edge of the support member for supporting the support member and for shielding the shroud member from a combustion gas flowing within the combustion turbine engine; and a layer of an abradable material disposed on a radially inner surface of the ceramic matrix composite shroud member for abradable wear against a rotating blade tip of the combustion turbine engine; the layer of abradable material and the ceramic matrix composite shroud member providing a degree of thermal insulation sufficient to maintain the metal support member below a predetermined temperature at respective points of direct contact between the ceramic matrix composite shroud member and the metal support member during operation of the combustion turbine engine. The blade shroud assembly may further include: a radially inner surface of the support member and a radially outer surface of the shroud member having respective closest points separated by a gap having a predetermined dimension; wherein a predetermined maximum dimension of the gap is selected so that a predetermined level of stress in the shroud member is not exceeded when the radially outer surface of the shroud member is deflected radially outwardly by the rotating blade tip to make contact with the radially inner surface of the support member.
A shroud assembly for sealing a cavity extending radially outward from a rotating blade tip to a blade ring of a combustion turbine engine to isolate the cavity from a combustion gas flowing past the blade tip is describe herein as including: a ceramic matrix composite member comprising a radially inner surface for wearing contact with the rotating blade tip and defining a primary pressure boundary for the combustion gas, the ceramic matrix composite member further comprising an arcuate portion defining a slot; a metal support member attached to a blade ring of the combustion turbine engine and comprising a radially outer surface separated from the radially inner surface by a gap and further comprising a portion extending into the slot for supporting the ceramic matrix composite member within the combustion turbine engine, the radially inner surface defining a secondary pressure boundary for the combustion gas in the event of failure of the ceramic matrix composite member; and the gap having a dimension sufficiently small to limit resonance of fluid surrounding the rotating blade tip in the event of failure of the ceramic matrix composite member. The gap may have a maximum dimension selected to control a level of stress developed in the ceramic matrix composite member when the ceramic matrix composite member is impacted by the rotating blade tip. The metal support member is selected to provide a predetermined resistance to further deflection of the ceramic matrix composite member when the ceramic matrix composite member is deflected to reduce the gap to zero.